System and method for reducing specific fuel consumption (sfc) in a turbine powered aircraft

ABSTRACT

A system for providing auxiliary power in an aircraft. A propulsion core comprises a compressor, a combustor, a turbine, and a shaft. An accessory unit comprises an accessory combustor, an accessory turbine, and an accessory shaft. A tank is configured to hold high pressure air and operably connected to the accessory unit. An electric generator comprises an electrical output and a mechanical input, with the mechanical input operably connected to the accessory shaft and the electrical output operably connected to an electric motor operably connected to the shaft. The electrical output is operably connected to an auxiliary power consuming device in the aircraft.

CROSS REFERENCES

This application is related to and incorporates by reference U.S. PatentPublication No. 2014/0248121.

BACKGROUND

As aircraft systems continue to develop more focus is on reducing bothemissions and fuel consumption, in conjunction with the goal of reducingoperational and capital costs. These goals may be achieved by increasingefficiencies of systems and/or optimizing systems for a given mission.The disclosed subject matter addresses these needs with a system thatallows optimization of accessory power unit sizing, operation and energyrecovery in conjunction with low cost power augmentation in the form ofonboard high pressure air.

Typically, prior art aircraft employ one or more primary power systemsresponsible for propulsion, the primary power systems also powersubsystems or accessory devices, such as the electrical, hydraulic,environmental, navigation and control systems. Prior art aircraft alsoinclude an auxiliary power unit (APU) for supplying power to theaccessory systems when the primary power systems are not available. TheAPU's are used when the aircraft is on the ground or when the aircraftis operating at lower speeds or altitudes, at cruise the primary powersystems provide all the power required by accessory systems and the APUis dead weight.

Generation of power for accessory systems by the primary power systemsare also complicated by the variable speeds which the primary powersystems operate over the course of a mission. This variation requiresadditional power conditioning equipment to ensure usable electricalpower to the accessory systems, which increases both capital andoperational costs. In aircrafts, every pound in equipment is a poundlost in payload, which corresponds to lost performance or profit.

SUMMARY

Embodiments of the present subject matter are presented herein as uniqueaircraft power systems. Additional embodiments include apparatuses,systems, devices, hardware, methods, and combinations for reducingemissions, fuel consumption, operational and capital costs. Furtherembodiments, forms, features, aspects, benefits, and advantages of thepresent application shall become apparent from the description andfigures provided herewith.

According to an aspect of the present disclosure, a system for providingauxiliary power in an aircraft comprises a propulsion core, an accessoryunit, a tank, and an electric generator. The propulsion core comprises acompressor, a combustor, a turbine, and a shaft. The accessory unitcomprises an accessory combustor, an accessory turbine, and an accessoryshaft. The tank is configured to hold high pressure air and operablyconnected to a high pressure air supply line between the tank and theaccessory unit. The electric generator comprises an electrical outputand a mechanical input, the mechanical input operably connected to theaccessory shaft and the electrical output operably connected to anelectric motor operably connected to the shaft and the electrical outputoperably connected to an auxiliary power consuming device in theaircraft.

In some embodiments an exhaust of the propulsion core is in thermalcontact with the high pressure air between the tank and the accessoryunit. In some embodiments the system further comprises a heat exchangercreating thermal contact with the exhaust and the high pressure air. Insome embodiments the high pressure air supply line provides highpressure air to the accessory combustor. In some embodiments the systemfurther comprises a fluid passage between the accessory unit and thecombustor, and the fluid passage is configured to provide accessory unitexhaust into the combustor.

In some embodiments the shaft operably powers a propulsion fan, lift fanor propeller. In some embodiments the system further comprises a highpressure air controller between the high pressure air tank and thecombustor. In some embodiments the system further comprises anelectrical controller between the electrical output and the electricmotor. In some embodiments the accessory unit further comprises anaccessory compressor.

According to another aspect of the present disclosure, a method of powermanagement in an aircraft is disclosed. The aircraft has a propulsioncore, an accessory turbine, and an accessory load. The method comprisesdetermining a maximum accessory load requirement of the aircraft;driving a generator with the accessory turbine to generate an electricaloutput as a function of the maximum accessory load requirement;maintaining the electrical output of the generator substantiallyconstant; and dividing the electrical output of the generator betweenthe accessory load and an auxiliary motor operably connected to a shaftof the propulsion core. The electrical output to the auxiliary motor isa function at least of the maximum accessory load requirement and aninstantaneous accessory load demand.

In some embodiments the method further comprises injecting high pressureair from an onboard high pressure air tank into an accessory combustorof the accessory turbine. In some embodiments the method furthercomprises pre heating the high pressure air with exhaust from thepropulsion core prior to injection into the combustor of the accessoryturbine. In some embodiments the method further comprises injecting anexhaust of the accessory turbine into a combustor of the propulsioncore.

In some embodiments the method further comprises charging the highpressure air tank from a source external to the aircraft. In someembodiments the method further comprises regulating the injection ofhigh pressure air and/or fuel into the accessory combustor to maintainthe electrical output of the generator. In some embodiments the methodfurther comprises sizing the generator and accessory turbine as afunction of the maximum accessory load requirement.

According to yet another aspect of the present disclosure, a method forreducing the specific fuel consumption for an aircraft mission isdisclosed. The method comprises predetermining characteristics of theaircraft mission; injecting high pressure air from an onboard tank intoa combustor of a power turbine; and controlling the rate of injection ofthe high pressure air into the combustor of the power turbine asfunction at least of the mass of the high pressure air in the onboardtank and predetermined characteristics of the mission.

In some embodiments the method further comprises heating the highpressure air with exhaust from the aircraft's primary propulsion systemprior to injection into the combustor and injecting exhaust from thepower turbine into a core combustor of the aircraft's primary propulsionsystem. In some embodiments the method further comprises determining anexpected duration of the mission; determining the amount of highpressure air available in the onboard tank, and determining a dischargemass flow rate of the high pressure air such that exhaustion of the highpressure air substantially corresponds to the end of the mission.

In some embodiments the method further comprises regulating the fuelsupply rate to the power turbine at least as a function of the dischargemass flow rate. In some embodiments the method further comprises furtherconverting work output of the power turbine into electricity andproviding electricity to the aircraft's auxiliary systems. In someembodiments the method further comprises driving the aircraft's primarypropulsion system with the electricity not utilized by the auxiliarysystems.

BRIEF DESCRIPTION OF THE FIGURES

FIG. 1 depicts an aircraft with a power system according to someembodiments of the disclosed subject matter.

FIG. 2 depicts an accessory power unit and high pressure air tankaccording to embodiments of the disclosed subject matter.

FIG. 3 depicts an arrangement of a primary power system, an accessorypower unit and a high pressure air tank according to an embodiment ofthe disclosed subject matter.

FIG. 4 depicts a flow chart for managing power from the generator of anaccessory power unit according to some embodiment of the disclosedsubject matter.

FIG. 5 illustrates a flow chart for reducing SFC based on a missionaccording to an embodiment of the disclosed subject matter.

FIG. 6 illustrates the recharge of high pressure air tanks according toan embodiment of the disclosed subject matter.

DETAILED DESCRIPTION

For the purposes of promoting an understanding of the principles of thesubject matter, reference will now be made to the embodimentsillustrated in the drawings and specific language will be used todescribe the same. It will nevertheless be understood that no limitationof the scope of the subject matter is thereby intended. Any alterationsand further modifications in the described embodiments, and any furtherapplications of the principles of the subject matter as described hereinare contemplated as would normally occur to one skilled in the art towhich the subject matter relates.

With reference to FIG. 1, an aircraft 1 is shown with gas turbine 10 asa primary power source used to provide propulsive power to the aircraft1 as to achieve or maintain a flight condition. While the primary powersource may take many forms and may include multiple engines, the primarypower source is preferably a gas turbine, and thus the subject matterwill be discussed with respect to such engine.

The illustrative embodiment in FIG. 1 depicts aircraft engine 10 as agas turbine engine and is shown including a compressor 12 forcompressing air, a combustor 14 for burning a mixture of fuel and thecompressed air, and a turbine 16 used to expand the combusted mixture offuel and air. The turbine 10 has a shaft 18 and an electrical drivemotor 19. Though the gas turbine engine 10 is shown as a single spoolturbojet engine, other embodiments can include additional numbers ofspools and can take other forms such as turbofan, turboprop, orturboshaft. In some embodiments the gas turbine engine 10 can be anadaptive cycle and/or variable cycle engine. It is contemplated that theengine 10 can have other variations and forms other than the few listedabove.

The aircraft 1 as shown in FIG. 1 also includes an accessory power unit20 and a high pressure air tank 21. The accessory power unit 20 and highpressure air tank 21 are shown in FIG. 2. The accessory power unit 20has a combustor (burner) 24, turbine 26 and output shaft 28, in additionto an optional compressor 22 that may be incorporated. The output shaftas shown is connected to an electrical generator 30. The accessory power20 unit provides power to the aircraft's non-propulsion subsystems 31 aswell as in some modes the propulsion subsystems. The high pressure airtank 21 is connected to the combustor 24 of the accessory power unit 20via supply lines 32. In addition, a valve 34 for controlling thedischarge of the high pressure air and a heat exchanger 36 are in thepath between the tank and the combustor 24. The heat exchanger 36pre-heats the high pressure air prior to injection. While FIG. 2 showsonly one high pressure air tank 21, the high pressure air tank 21 maycomprise multiple tanks connected with multiple supply lines.

FIG. 3 is a diagram of the interaction between the primary propulsionsource, turbine engine 10, the accessory power unit 20 and high pressureair tank 21. The output of both the turbine engine 10 and the accessorypower unit 20 include both useable power and heat. Utilizing an opentopping cycle like arrangement, exhaust from the turbine engine 10 isfed through passage 40 into the heat exchanger 36 to pre-heat the highpressure air prior to injection into the accessory power unit 20.Likewise the exhaust from the accessory power unit 20 is directed intothe combustor 14 of turbine engine 10 via passage 42. In this manner thedisclosed arrangement recovers some of the energy previously lost. U.S.Patent Publication No. 2014/0223918 and U.S. Pat. No. 9,003,763,incorporated herein by reference, describe the thermodynamic gains in aterrestrial system with respect to a topping cycle that are applicableto an open system as described herein.

As illustrated in FIG. 3 the accessory power unit 20 in addition tooutputting heat energy also through the generator 30 outputs electricalpower which is supplied to the aircraft's subsystems 31 via supply line44 and is supplied to the drive motor 19 via supply line 46. Thegenerator 30 further advantageously includes an electrical (power)controller (not shown) for distributing the power between the subsystems31 and drive motor 19.

FIG. 3 also shows the operable relationship described above between thehigh pressure air tank 21, heat exchanger 36 and the accessory powerunit 20. The heat exchanger 36 may be direct in which case exhaust heatis directly transferred to the high pressure air, or indirect in which asecondary fluid is used to exchange energy between the two. While notshown in FIG. 1 or 3, the shaft 18 drives a propeller, fan, rotor, orpump to provide propulsion to the aircraft 1, or drives a generator orpump which in turn drives a propeller, fan, rotor or pump.

FIG. 4 is illustrative of a flow chart for managing the power in theaircraft 1. The maximum accessory load requirement of the aircraft ifnot known for a given aircraft 1 is determined, upon which the accessorypower unit 20 may be sized as shown in Block 401. The load requirementmay be a function of the maximum load of each of the non-propulsivesubsystems 31 of the aircraft 1 plus a safety factor, or may be refinedto be a function of the maximum load of all subsystems that would beconcurrently in operations plus a safety factor. This maximum loadrequirement is preferably a function of the aircraft and not of aparticular mission, as the sizing and optimum state of the accessorypower unit 20 particular to the aircraft is based upon this requirementduring the design or retrofitting stage.

The accessory power unit 20 during operation of the aircraft 1 drivesthe generator 30 to create a constant electrical power output as shownin Block 403. The generator 30, along with the accessory power unit 20,are advantageously matched to operate at a set optimum speed and poweroutput. The generator 30 may be directly driven by the accessory powerunit 20 or via a gearbox. Preferably the combination of accessory powerunit 20 and generator 30 run at a constant rpm selected as a function ofthe set's efficiency.

The constant electrical output from the generator 30 is distributedbetween the subsystem load 31 and the auxiliary electric motor 19 whichis connected to the shaft 18 of the propulsion core (turbine engine 10)as shown in Block 405. The generator set (accessory power unit 20 andgenerator 30) runs constantly at its most efficient power settingindependent of the instantaneous demand from the non-propulsionsubsystems 31, resulting in reductions in SFC of around 2%. Electricenergy produced and not consumed by the non-propulsion subsystems 31 isdistributed to the propulsion core via the drive motor 19, thus reducingthe fuel requirements of the turbine engine 10. The constant input andoutput from the generator 30 eliminates need for a separateintermittently used APU and power conversions from rpm dependent outputfrequencies/voltages from prior art propulsion core generators.

As shown in Block 407, high pressure air is injected into the combustor24 of the accessory power unit 20 and preferably the injected air ispreheated by the exhaust from the turbine engine 10. The introduction ofheated pressurized air into the combustor reduces the amount of workextracted from the accessory turbine 26 by the compressor 22 (ifpresent) and reduces the SFC of the accessory power unit 20 whilemaintaining the power output of the generator set in comparison tooperations in which high pressure and heated high pressure injection asdescribed herein is not utilized. The injection of high pressure air canresult in a 5-15% reduction in SFC for the accessory power unit 20. Theinjection of high pressure air into the auxiliary power unit 20 may alsobe controlled to adjust for operating conditions which effect the SFC ofthe auxiliary power unit 20 such as, humidity, temperature, pressure,flow disruptions, oxygen depletion or airborne particles.

Injecting the exhaust from the accessory power unit 20 into thecombustor 14 of the turbine engine 10 as shown in Block 409 may alsoincrease SFC of the turbine engine 10 approximately 2% depending on therelative size of the accessory power unit 20 and the turbine engine 10.

As illustrated in FIG. 5, the SFC of an aircraft can advantageously bereduced by controlling the injection of the high pressure air as afunction of the mission (flight plan). Characteristics of the aircraftmission are predetermined as shown in Block 501. The predominatedcharacteristic of the mission being duration, however othercharacteristics may include required reserve, contingent operations,subsystem loads, demand surges. From the onboard tank 21, high pressureair is injected into the accessory power unit 20, as shown in Block 503and controls the air as a function shown in Block 505. The rate ofinjection (discharge rate) of the high pressure air into the combustor24 of the power turbine is preferably controlled as function at least ofthe mass (or capacity) of the high pressure air in the onboard tank andat least one predetermined characteristic of the mission.

Determining the mission characteristics further includes: determining anexpected duration of the mission, determining the amount of highpressure air is available, and determining the injection pressure of thehigh pressure air. It is preferable that in determining a discharge massflow rate of the high pressure air based on the mission characteristics,the exhaustion of the high pressure air substantially corresponds to theend of the mission, exclusive of reserve air. High pressure air not usedor held in reserves amounts to a lost opportunity for SFC savings inboth the auxiliary power unit 20 and the turbine engine 10, as well asincreased carried dead weight.

As discussed previously, the exhaust from the accessory power unit 20may advantageously be injected into the combustor of the turbine engine10 as well as heating the high pressure air with exhaust from theaircraft's primary propulsion system prior to injection into thecombustor. In response to the high pressure air injection, the fuelsupply rate to the accessory power unit 20 may be reduced at least as afunction of the discharge mass flow rate. Alternatively, or inconjunction, if a compressor 22 is included in the auxiliary power unit20, the mass flow rate of air provided by the compressor 22 of theaccessory power unit 20 may also be reduced as a function of thedischarge mass flow rate. Likewise, the fuel to the primary propulsionunit 10 may also be reduced in response to the high pressure airinjection.

As described above, the work output of the accessory power unit 20 isconverted into electricity and provided the aircraft's non-propulsionsubsystems 31. The electricity not utilized by these subsystems 31 isutilized to drive the aircraft's primary propulsion system (turbineengine 10) via electric motor 19.

While not shown, a dynamic assessment and control of the air dischargerate may also be advantageous. For example, a plane on standby on a hottarmac may wish to increase the discharge rate to avoid ingesting alarge amount of low pressure hot gas into the compressor 22 (if present)which may be detrimental to the efficiency of the accessory power unit20. Another example of dynamic reassessment may be at the end of amission, if at which all the air has not been discharged, increasing thedischarge rate to ensure it is all used prior to landing and provideexcess power to the turbine engine 10 via motor 19. As noted above,ground equipment may quickly and inexpensively recharge the highpressure tank 21 thus obviating the need to continue to maintain areserve mass or include the optional compressor 22.

Advantageously the discharge valve 34 is automatically controlled by acontroller. The controller can include one or more Arithmetic LogicUnits (ALUs), Central Processing Units (CPUs), memories, limiters,conditioners, filters, format converters, or the like which are notshown to preserve clarity. In one form, the controller is of aprogrammable variety that executes algorithms and processes data inaccordance with operating logic that is defined by programminginstructions (such as software or firmware). Alternatively oradditionally, operating logic for the controller can be at leastpartially defined by hardwired logic or other hardware. In oneparticular form, the controller is configured to operate as a FullAuthority Digital Engine Control (FADEC); however, in other embodimentsit may be organized/configured in a different manner as would occur tothose skilled in the art. It should be appreciated that the controllercan be exclusively dedicated to controlling operation of the dischargevalve, or may additionally and/or alternatively be used in theregulation/control/activation of one or more other subsystems or aspectsof the aircraft 1.

The high pressure air tank 21 is capable of being pressurized to avariety of pressures and can be any size and/or shape and have a varietyof constructions. More than one tank 21 can be provided in anyembodiment. The high pressure air tank 21 is capable of being chargedwith pressurized air while installed on the aircraft 1 or can be removedfor servicing from the aircraft 1. In some forms the high pressure airtank 21 can be charged using one or more of the aircraft engines;however, it is contemplated that the high pressure air tank 21 isrecharged using another source external to the aircraft 1. The highpressure air tank 21 may also be recharged in flight in some modes ofoperation and may be recharged while the aircraft 1 is on the ground inothers.

As depicted in the illustrative embodiment, a controllable valve 34(discharge valve) is disposed between the high pressure air in the highpressure air tank 21 and the gas turbine engine 10. An air supply line32 can be disposed between the high pressure air tank 21 and the valve34, as well as another air supply line 32 between the valve 34 and thegas turbine engine 10. In some embodiments, the valve 34 may beconnected with the tank 21 such that an intermediate air line is notneeded. In some embodiments multiple air lines can be connected betweenmultiple valves 34 to couple the high pressure air tank 21 to theaccessory power unit 20. The multiple valves 34 can be connecteddirectly to the high pressure air tank 21 or can be connected withmultiple air lines. In still other embodiments, more than one air tank21 can be provided in the aircraft 1.

The valve 34 can take on a variety of forms and can be actuated using avariety of techniques. To set forth just a few examples, the valve 34can be driven or powered by devices that are mechanical, hydraulic,manual, electrical, electromechanical, or combinations thereof. Thevalve 34 can be arranged to have only two positions, open or closed, orcan be a valve that provides any number of intermediate positions. Inaddition, the valve 34 can be capable of being commanded to any givenposition at a common rate in some embodiments and a variety of rates inothers. Any variety of flow rates of the pressurized air can be providedthrough the valve 34. The valve 34 can be a one-time use valve or can beactuated a number of times to different positions. To set forth just afew further examples of variations, the valve 34 can be a ball valve,butterfly valve, check valve, gate valve, needle valve, piston valve,spool valve, or a poppet valve. In some forms the valve 34 can act as apressure regulator. More than one valve 34 can be provided to admitpressurized air from the high pressure air tank 21 to the accessorypower unit 20, in which case the valves 34 can be, but need not be, thesame. In embodiments of the aircraft 1 in which multiple valves 34 areused with multiple air tanks 21, not all valves 34 need be the same.

The valve 34 can be controlled by the controller. In one embodiment thecontroller is capable of providing a signal to open the valve 34. Thecontroller can be used in some embodiments to control the rate at whichthe valve 34 is opened and/or the position to which the valve 34 isopened and thus the discharge mass flow rate of the high pressure air.In some embodiments the controller is capable of controlling the valve34 at any variety of positions between an open and closed position. Thecontroller can provide a signal to open and/or close the valve 34 basedupon a request received from an operator, such as through a switchlocated in a cockpit, to set forth just one non-limiting example. Instill other embodiments a switch can directly command the valve 34without the need of the controller. In some embodiments, the controllercan monitor aircraft systems and depending on a control algorithmprovide signals to the valve 34. Other situations could also give riseto the valve 34 opening and admitting pressurized air into the accessorypower unit 20.

The high pressure air tank 21 can be used to provide an increase in massflow in the auxiliary power unit, above that provided by an optionalcompressor 22 if present. Consequently, when the mass flow increases thefuel flow can be decrease to maintain the same power output.

The controller can automatically operate the valve 34 prior to commandand/or confirmation from the pilot. In these situations, an alert can beprovided to the cockpit to notify the crew of the automatic engagementof the system.

Turning now to FIG. 6, one form of the present application is shown inwhich the high pressure air tank 21 can be recharged. The pump 602 cantake any variety of forms such as centrifugal pumps, axial pumps, screwpumps, gear pumps, lobe pumps, vane pumps, piston pumps, diaphragmpumps, and plunger type pumps, to set forth just a few non-limitingexamples. The pump 602 can be mechanically driven in some embodiments,but in others the pump 602 can be a hydraulic, electrical, or pneumaticdriven pump. More than one pump 602 can be used for the high pressureair tank 21. A pump 602 can be used to charge more than one tank 21, inwhich case some embodiments may include a valve 604 to select betweenthe tanks 21. The engine 606 is preferably ground based and external tothe aircraft, such as another aircraft or ground servicing equipment. Itis also advantageous that connection valves 608 by which the charginglines 609 are connected are a type of quick connect valves to ensuresimple and rapid recharging of the tank 21.

In the illustrative embodiment the primary propulsion core (turbineengine 10) drives an external load which can represent a propeller, fan,pump or a rotor, to set forth some non-limiting examples. Someembodiments need not include an external load in that the accelerationof the air itself through the engine provides the propulsive thrust.

An aspect of the disclosed subject matter as noted previously is theelimination of the compressor 22 for the auxiliary power unit 20 infavor of the high pressure injection. Alternatively, the compressor 22of the auxiliary power unit 20 may be sized to merely augment the massflow rate provided by the high pressure air injection or serve in onlyan emergency situation, in which case it would be designed to deliveronly a minimum mass flow. Either the compressor entirely eliminated orwith a reduced sized compressor, weight and cost are advantageouslyfurther reduced.

As used above, the term “aircraft” includes, but is not limited to,helicopters, airplanes, unmanned space vehicles, fixed wing vehicles,variable wing vehicles, rotary wing vehicles, unmanned combat aerialvehicles, tailless aircraft, hover crafts, and other airborne and/orextraterrestrial (spacecraft) vehicles. Similarly, the term “air” can beany suitable fluid which may or may not have the elemental compositionof air. It should also be noted that the use of propulsion core,propulsion system and turbine engine are used synonymously throughout.Likewise non-propulsion subsystems, accessory subsystem and accessoryloads are also used synonymously.

Although examples are illustrated and described herein, embodiments arenevertheless not limited to the details shown, since variousmodifications and structural changes may be made therein by those ofordinary skill within the scope and range of equivalents of the claims.

What I claim is:
 1. In an aircraft, a system for providing auxiliarypower comprising: a propulsion core, the propulsion core comprising acompressor, a combustor; a turbine and a shaft; an accessory unit, theaccessory unit comprising an accessory combustor; an accessory turbineand an accessory shaft; a tank configured to hold high pressure air, thetank operably connected to a high pressure air supply line between thetank and the accessory unit; and, an electric generator comprising anelectrical output and a mechanical input, the mechanical input operablyconnected to the accessory shaft and the electrical output operablyconnected to an electric motor operably connected to the shaft and theelectrical output operably connected to an auxiliary power consumingdevice in the aircraft.
 2. The system of claim 1, wherein an exhaust ofthe propulsion core is in thermal contact with the high pressure airbetween the tank and the accessory unit.
 3. The system of claim 2,further comprising a heat exchanger creating thermal contact with theexhaust and the high pressure air.
 4. The system of claim 1, wherein thehigh pressure air supply line provides high pressure air to theaccessory combustor.
 5. The system of claim 1, further comprising afluid passage between the accessory unit and the combustor, the fluidpassage configured to provide accessory unit exhaust into the combustor.6. The system of claim 1, wherein the shaft operably powers a propulsionfan, lift fan or propeller.
 7. The system of claim 1, further comprisinga high pressure air controller between the high pressure air tank andthe combustor.
 8. The system of claim 1, further comprising anelectrical controller between the electrical output and the electricmotor.
 9. The system of claim 1, wherein the accessory unit furthercomprises an accessory compressor.
 10. A method of power management inan aircraft, the aircraft having a propulsion core, an accessory turbineand an accessory load, the method comprising: determining a maximumaccessory load requirement of the aircraft; driving a generator with theaccessory turbine to generate an electrical output as a function of themaximum accessory load requirement; maintaining the electrical output ofthe generator substantially constant; and, dividing the electricaloutput of the generator between the accessory load and an auxiliarymotor operably connected to a shaft of the propulsion core; wherein theelectrical output to the auxiliary motor is a function at least of themaximum accessory load requirement and an instantaneous accessory loaddemand.
 11. The method of claim 10, further comprising injecting highpressure air from an onboard high pressure air tank into an accessorycombustor of the accessory turbine.
 12. The method of claim 11, furthercomprising pre heating the high pressure air with exhaust from thepropulsion core prior to injection into a combustor of the accessoryturbine.
 13. The method of claim 10, further comprising injecting anexhaust of the accessory turbine into a combustor of the propulsioncore.
 14. The method of claim 11, further comprising charging the highpressure air tank from a source external to the aircraft.
 15. The methodof claim 11, further comprising regulating the injection of highpressure air and/or fuel into the accessory combustor to maintain theelectrical output of the generator.
 16. The method of claim 11, furthercomprising sizing the generator and accessory turbine as a function ofthe maximum accessory load requirement.
 17. A method for reducing thespecific fuel consumption for an aircraft mission, the methodcomprising: predetermining characteristics of the aircraft mission;injecting high pressure air from an onboard tank into a combustor of apower turbine; and, controlling the rate of injection of the highpressure air into the combustor of the power turbine as function atleast of the mass of the high pressure air in the onboard tank andpredetermined characteristics of the mission.
 18. The method of claim 17further comprising heating the high pressure air with exhaust from theaircraft's primary propulsion system prior to injection into thecombustor and injecting exhaust from the power turbine into a corecombustor of the aircraft's primary propulsion system.
 19. The method ofclaim 17, further comprising determining an expected duration of themission; determining the amount of high pressure air available in theonboard tank, and determining a discharge mass flow rate of the highpressure air such that exhaustion of the high pressure air substantiallycorresponds to the end of the mission.
 20. The method of claim 19,further comprising regulating the fuel supply rate to the power turbineat least as a function of the discharge mass flow rate.